System For Reducing Combustion Noise And Improving Cooling

ABSTRACT

A novel and improved system for cooling and reducing combustion noise in a gas turbine combustor is disclosed. The system includes a flow sleeve for a gas turbine combustor comprising a tubular portion and a conical portion, with a plurality of flow straighteners extending radially inward and between the tubular and conical portions and a plurality of rows of cooling holes extending about the flow sleeve, and an aft ring secured to the outlet end of the conical portion, where the aft ring includes an overlapping piston ring that is able to expand or contract in diameter. A combustion liner extends through the flow sleeve and engages an inlet of a transition duct while the piston ring of the flow sleeve also engages the transition duct to form a seal.

CROSS-REFERENCE TO RELATED APPLICATIONS

None.

TECHNICAL FIELD

The present invention generally relates to a system for reducingcombustion noise and directing cooling air into a gas turbine combustor.

BACKGROUND OF THE INVENTION

In a typical gas turbine engine used in a powerplant application, aplurality of combustors are arranged in an annular array about acenterline of the engine. The combustors receive pressurized air fromthe engine's compressor, add fuel to create a fuel/air mixture, andignite the mixture to produce hot combustion gases. The hot combustiongases exit the combustors and enter a turbine, where the expanding gasesare utilized to drive a turbine, which is in turn coupled through ashaft to the compressor. The engine shaft is also coupled to a shaftthat drives a generator for generating electricity.

The combustors typically include at least a pressurized case and acombustion liner contained within the case. The fuel, which is suppliedby a plurality of fuel nozzles, mixes with air and reacts (i.e. ignites)within the combustion liner. In order to actively cool the combustionliner, the compressed air that is used for combustion is first directedthrough the pressurized case and along the combustion liner. The airthen mixes with fuel and reacts in the combustion liner.

Prior art configurations of combustors include a flow sleeve extendingthrough the case and used to support a combustion liner in place withinthe combustor. The flow sleeve often includes a series of holes throughwhich compressed air passes. Air passing through these holes is intendedto impinge on the combustion liner wall. However, in prior gas turbinecombustor configurations, air streams have been known to be ineffectivein maintaining active cooling through impingement, thereby leading topremature degradation and damage of the combustion liner.

SUMMARY

In accordance with the present invention, there is provided a novel andimproved system for cooling and reducing combustion noise in a gasturbine combustor. An embodiment of the present invention includes aflow sleeve for a gas turbine combustor comprising a tubular portion anda conical portion, with a plurality of flow straighteners extendingradially inward and between the tubular and conical portions. The flowsleeve also includes a plurality of rows of cooling holes extendingabout the flow sleeve and an aft ring secured to the outlet end of theconical portion, where the aft ring includes a receptacle containing apiston ring that is able to expand or contract in diameter.

In another embodiment of the present invention, a cooling system for agas turbine combustor is disclosed comprising a flow sleeve, acombustion liner, and a transition duct. The flow sleeve includes atubular portion, a conical portion, a plurality of rows of coolingholes, and an aft ring having a receptacle and a piston ring containedwithin the receptacle. A cooling passage is formed between thecombustion liner and the flow sleeve and directs air received from theflow sleeve to the inlet of the combustor. The transition duct iscoupled to the combustion liner at the liner outlet. A piston ring ispositioned in the aft end of the flow sleeve to create a seal betweenthe flow sleeve such that compressor air is only permitted to enter thepassageway through holes in the flow sleeve.

In yet another embodiment of the present invention, a sealing device foran aft end of a gas turbine combustor, the device comprising a sleevehaving a conical portion and an aft ring with a receptacle. A pistonring is located in the receptacle, where the piston ring is able toexpand or contract in order to form a seal with the transition duct.

Additional advantages and features of the present invention will be setforth in part in a description which follows, and in part will becomeapparent to those skilled in the art upon examination of the following,or may be learned from practice of the invention. The instant inventionwill now be described with particular reference to the accompanyingdrawings.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

The present invention is described in detail below with reference to theattached drawing figures, wherein:

FIG. 1 is a cross section view of a portion of a gas turbine combustorin accordance with the prior art;

FIG. 2 is a detailed cross section view of a portion of the gas turbinecombustor of FIG. 1 in accordance with the prior art;

FIG. 3 is a cross section view of a portion of a gas turbine combustorin accordance with an embodiment of the present invention;

FIG. 4 is a detailed cross section view of a portion of the gas turbinecombustor of FIG. 3 in accordance with an embodiment of the presentinvention; and,

FIG. 5 is a cross section view of a flow sleeve depicted in FIG. 3 inaccordance with an embodiment of the present invention.

DETAILED DESCRIPTION

The subject matter of the present invention is described withspecificity herein to meet statutory requirements. However, thedescription itself is not intended to limit the scope of this patent.Rather, the inventors have contemplated that the claimed subject mattermight also be embodied in other ways, to include different components,combinations of components, steps, or combinations of steps similar tothe ones described in this document, in conjunction with other presentor future technologies.

The present invention is directed generally towards a system forreducing combustion noise controlling the flow of cooling air to acombustion liner. Referring initially to FIGS. 1 and 2, a portion of agas turbine combustor according to the prior art is shown in crosssection. A flow sleeve 100 includes a combustion liner 102 containedtherein. The combustion liner 102 engages a double wall transition duct104 for purposes of passing the flow of hot combustion gases from thecombustion liner 102 to an inlet of a turbine (not shown). In the priorart configuration, there existed a piston ring in an annular gap 106 atthe aft end of the flow sleeve 100 and the inlet of transition duct 104.The piston ring serves to minimize the flow of air through the gap 106created by the flow sleeve 100 and transition duct 104. However, theamount of cooling air passing through the annular gap 106 is notminimized due to a large opening in the piston ring that is created atoperating temperatures, thereby affecting the even distribution of airflow to the head end of the combustor.

The present invention can be better understood when considering FIGS.3-5. Referring initially to FIG. 3, a cooling system for a gas turbinecombustor 200 is disclosed providing improved control of the flow ofcooling air thereby also providing improved control of combustion noisedue to variations in rates of air provided for mixing with fuel andresulting combustion.

Referring now to FIGS. 3-5, the gas turbine combustor 200 comprises aflow sleeve 202 having a tubular portion 204 with a tubular outlet 204Aand a tubular inlet 204B. The flow sleeve 202 also includes a conicalportion 206 that is connected to the tubular portion 204 at the tubularinlet 204B, where the conical portion has a conical outlet 206A and aconical inlet 206B. The flow sleeve 202 also includes a plurality ofrows of cooling holes 208, where each row 208 extends about a perimeterof the flow sleeve 202. An aft ring 210 is located about the conicalinlet 206B and has a receptacle 212 for receiving a piston ring 214. Thepiston ring 214 incorporates a split-ring design so as to be capable ofexpanding or contracting in diameter depending on the size of the matinghardware. As such, the piston ring 214 provides an overlap area (notshown) so as to maintain a flow blockage under all conditions. In anembodiment of the present invention, the flow sleeve 202 also includes aplurality of flow straighteners 216, as depicted in FIG. 3. The flowstraighteners 216 are designed to channel the flow of cooling air fromcooling holes 208 axially thereby reducing swirling characteristics ofthe cooling air flow.

A combustion liner 220 is located within the flow sleeve 202 and extendsthrough the axial length of the flow sleeve 202. The combustion liner220 is generally tubular in shape and, as a result of its locationwithin the combustion system, creates a cooling passage 222 between thecombustion liner 220 and flow sleeve 202. It is the air passing throughpassageway 222 that flow straighteners 216 attempt to straighten so asto achieve a more uniform flow path.

The present invention also includes a transition duct 230 which iscoupled to the outlet of the combustion liner 220 for receiving the hotcombustion gases from the combustion liner 220. Due to the geometricrequirements of the combustion liner 220 compared to the inlet region ofa turbine, the transition duct 230 transitions from a generally circularcross section at the duct inlet to a generally rectangular cross sectionat the duct outlet.

As discussed above, the flow sleeve 202 includes a piston ring 214. Thepiston ring 214 creates a seal between the flow sleeve 202 and thetransition duct 230. Because the piston ring is expandable, it has theability to adjust to various tolerance conditions and differentialthermal growth between the flow sleeve 202, which is relatively cold,and the transition duct 230, which is relatively hot, due to thecombustion gases contained within. By maintaining a constant contactbetween the piston ring 214 and the transition duct 230, a seal isformed that prevents compressor air from entering the passageway 222from the aft end of the flow sleeve 202. Instead all of the compressedair is directed towards the conical and tubular portions of the flowsleeve 202 so that it may enter through the plurality of rows of coolingholes 208.

As discussed above, the flow sleeve 202 includes a plurality of coolingholes 208 arranged in multiple rows. As shown in FIGS. 3 and 4, at leastone of the rows of cooling holes 208 is located in the tubular portion204 and at least one of the rows is located in the conical portion 206.In an embodiment, the cooling holes 208 are oriented generallyperpendicular with respect to the tubular portion 204 and conicalportion 206. In an alternate embodiment, the cooling holes 208 can beoriented at a surface angle a relative to both the tubular portion 204.The surface angle a can vary depending on the flow sleeve configuration,but is preferable between approximately 10 and 20 degrees. The coolingholes 208 can also vary in diameter between 0.5 inches and 1.75 inches.The hole sizes of the flow sleeve are specifically sized based on theirdistance to the combustion liner. This enables a more consistent flowaround the flow sleeve and liner annulus, which leads to a more evenflow to the head end of the combustor. The hole size is based generallyon the ratio of the hole area to the flow area from the hole to thecombustion liner.

In an embodiment of the invention, the flow sleeve 202 has four rows ofholes 208 with each row having 24 holes. The aft-most row of holes 208A,closest to the aft ring 210, have a diameter of approximately 0.95inches, with the adjacent row moving forward, indicated as 208B, have adiameter of approximately 1.5 inches, while the next row 208C have adiameter of approximately 1.575 inches, and the forward-most row ofholes 208D having a diameter of approximately 0.65 inches. As a resultof the cooling holes 208 being placed in the flow sleeve 202, thecooling holes 208 impart a jet of cool air onto the combustion liner 220that is located in the flow sleeve 202. The jet impinges air on theliner which cools the liner wall and then the air travels upstreamtowards an inlet to the combustion liner 220.

The present invention has been described in relation to particularembodiments, which are intended in all respects to be illustrativerather than restrictive. Alternative embodiments will become apparent tothose of ordinary skill in the art to which the present inventionpertains without departing from its scope.

From the foregoing, it will be seen that this invention is one welladapted to attain all the ends and objects set forth above, togetherwith other advantages which are obvious and inherent to the system andmethod. It will be understood that certain features and sub-combinationsare of utility and may be employed without reference to other featuresand sub-combinations. This is contemplated by and within the scope ofthe claims.

1. A flow sleeve for a gas turbine combustor comprising: a tubularportion having a tubular inlet and a tubular outlet; a conical portionhaving a conical inlet and a conical outlet, the conical inlet securedto the tubular outlet of the tubular portion; a plurality of flowstraighteners extending radially inward from the tubular portion and theconical portion; a plurality of rows of cooling holes, each row ofcooling holes extending about the flow sleeve; an aft ring having areceptacle, the aft ring secured to the conical outlet; and a pistonring positioned within the receptacle, the piston ring capable ofexpanding or contracting in diameter.
 2. The flow sleeve of claim 1,wherein at least one of the plurality of rows of cooling holes islocated in the tubular portion.
 3. The flow sleeve of claim 2, whereinat least one of the plurality of rows of cooling holes is located in theconical portion.
 4. The flow sleeve of claim 1, wherein the coolingholes impart a jet of cooling air onto a combustion liner located withinthe flow sleeve.
 5. The flow sleeve of claim 1, wherein the piston ringis able to slide axially and radially within the receptacle.
 6. The flowsleeve of claim 1, wherein the plurality of rows of holes are generallyperpendicular to the tubular portion and the conical portion.
 7. Theflow sleeve of claim 1, wherein the cooling holes range in diameter fromapproximately 0.5 inches to approximately 1.75 inches.
 8. The flowsleeve of claim 7, wherein the conical portion is oriented at an anglebetween 10 degrees and 20 degrees relative to the tubular portion.
 9. Acooling system for a gas turbine comprising: a flow sleeve comprising: atubular portion having a tubular inlet and a tubular outlet; a conicalportion having a conical inlet and a conical outlet, the conical portionconnected to the tubular portion at the tubular outlet; a plurality ofrows of cooling holes, each row extending about the perimeter of theflow sleeve; an aft ring located about the conical outlet and having areceptacle; and a piston ring positioned within the receptacle; acombustion liner comprising a generally tubular body and extendingthrough the flow sleeve thereby forming a cooling passage therebetween,the combustion liner having a liner inlet and a liner outlet; and atransition duct coupled to the liner outlet, the transition ductreceiving hot combustion gases from the combustion liner, the transitionduct having an outer wall transitioning from a generally circular ductinlet to a generally rectangular duct outlet; wherein the piston ring ofthe flow sleeve creates a seal between the flow sleeve and thetransition duct, thereby preventing compressor air from entering apassageway formed between the flow sleeve and the combustion liner, andinstead directing the compressed air to enter the passageway through theplurality of rows of cooling holes.
 10. The cooling system of claim 9,wherein at least one of the plurality of rows of cooling holes islocated in the tubular portion of the flow sleeve.
 11. The coolingsystem of claim 9, wherein at least one of the plurality of rows ofcooling holes is located in the conical portion of the flow sleeve. 12.The cooling system of claim 9, wherein the plurality of rows of coolingholes are oriented generally perpendicular to a surface of the tubularportion.
 13. The cooling system of claim 9, wherein the plurality ofrows of cooling holes are oriented at an angle relative to a surface ofthe tubular portion.
 14. The cooling system of claim 9, wherein theconical portion is oriented at an angle between 10 degrees and 20degrees relative to the tubular portion.
 15. A sealing device for an aftend of a gas turbine combustor comprising: an axially-extending sleevehaving a conical portion; an aft ring with a U-shaped receptacle at anoutlet end of the conical portion; and a piston ring located in theU-shaped receptacle, the piston ring capable of expanding or contractingin diameter upon placement of an inlet end of a transition duct into theoutlet end of the conical portion; wherein a seal is formed between thepiston ring and the transition duct, thereby preventing air fromentering a combustor through a gap between the sleeve and the transitionduct.
 16. The sealing device of claim 15, wherein the U-shapedreceptacle has a width greater than a width of the piston ring.
 17. Thesealing device of claim 15, wherein the U-shaped receptacle has adiameter greater than a diameter of the piston ring.
 18. The sealingdevice of claim 15, wherein the piston ring has a generally rectangularcross section.